1. Field of the Invention
This invention relates to the field of transonic aerodynamics and, more particularly, to a splitter plate extending from the blunt trailing edge of a wing for a transonic airplane.
2. Description of the Prior Art
The aerodynamic drag of modern transonic wings is comprised of two components, parasitic drag and compressibility drag, which is also known as wave drag. Parasitic drag can further be divided into the components of friction, form drag, and base drag. At lower Mach numbers, only parasitic drag is present. As the Mach number is increased to the regime above 0.6, the fluid flow across the wing's upper surface becomes supersonic and results in the formation of shock waves. The drag caused by the irreversible compression associated with the formation of these shock waves is called compressibility drag.
Compressibility drag increases dramatically with the strength of the formed shock waves. Modern wings use both wing sweep and airfoil shaping to delay the onset of shock wave formation over the top of the wing until the freestream velocity is as high as possible, while at the same time maintaining high lift and low parasitic drag.
While preserving desirable lift and drag characteristics, it is advantageous to have as thick a wing as possible for several reasons. Firstly, a thicker wing is able to save weight by using deep, structurally efficient beams which are light but bulky. By contrast, a relatively thin wing must rely for its strength on material thickness which, despite a thinner profile, weighs more than the thicker wing which obtains the same strength through the use of deep beams. Further, as the wings on commercial transport airplanes are used to store fuel, a thicker wing provides for increased range. A thicker wing also facilitates storage of landing gear therein. The inherent problem is that a thicker wing promotes higher airstream velocities across both its top and bottom surfaces relative to a thinner wing, and thus can cause the onset of compressibility drag at a freestream velocity below the design cruise velocity of the airplane.
In U.S. Pat. No. 3,952,971, Richard T. Whitcomb shows what is called a supercritical airfoil having an upper surface contoured to control flow acceleration and pressure distribution thereon in order to prevent or mitigate shock wave formation on the upper surface. The trailing edge section is more highly cambered than usual to improve overall lifting efficiency. Use of this design allows a wing to be thickened without inducing compressibility drag.
Unfortunately, highly aft-cambered airfoils such as the supercritical airfoil tend to be thin in the region of the trailing edge flap. This thinness makes it difficult to design trailing edge flaps. Further, adverse viscous boundary layer effects have been found to be more significant for highly aft-loaded airfoils. As a result, an appreciable amount of the aft camber is effectively lost due to viscous boundary layer decambering on the upper surface near the trailing edge and in the cove region of the lower surface. Thus, the full theoretical benefit of the supercritical airfoil has not been realized in practice.
An advancement over the supercritical airfoil is shown in U.S. Pat. No. 4,858,852, issued to Preston A. Henne et al. The foregoing patent discloses an airfoil having a trailing edge with diverging upper and lower surfaces. This design has been found to delay the onset of compressibility drag and thus provide the designer with the option of increasing the thickness of the airfoil.
Another approach to the problem of compressibility drag and the related variable of airfoil thickness is shown in U.S. Pat. No. 4,542,868 issued to James A. Boyd. The foregoing reference shows a wedge affixed to the bottom surface of the trailing edge of an airfoil.
The mutual problem inherent to the approaches of Henne et al and Boyd is that they result in the airfoil having a blunt trailing edge. The increased pressure drag, also known as base drag, which typically accompanies the use of a blunt trailing edge thus stands to offset the gains otherwise afforded by the aforementioned improvements in airfoil design.
Tests have shown that a splitter plate extending normally from the geometric middle of the blunt trailing edge of an axisymmetrical airfoil having a zero angle of attack (no lift) will significantly reduce the base drag. A discussion and analysis of the results of such a series of tests is provided by E. Saltzman and J. Hintz, Flight Evaluation of Splitter-Plate Effectiveness in Reducing Base Drag at Mach Numbers from 0.65 to 0.90, National Aeronautics and Space Administration Technical Memorandum X-1376 (May 1967). However, these tests were conducted on airfoils dissimilar to modern cambered airfoils which operate under lifting conditions. Moreover, the prior art tests typically used a blunt trailing edge having a thickness to mean chord length ratio several times greater than the ratio for blunt trailing edges used on the wings of commercial transport airplanes, thus rendering the applicability of their results to modern wings all the more conjectural.
Finally, the tests of the prior art used a splitter plate lying collinear with the chord line of an axisymmetrical airfoil. One versed in aerodynamics would expect that attaching a splitter plate to the blunt trailing edge of a cambered airfoil in the aforementioned collinear alignment with the chord line would effectively decamber the airfoil and precipitate the onset of compressibility drag at a lower freestream velocity than would be the case in the absence of the splitter plate.
In summary, the significant design advantages attendant to using thicker wings can presently be realized by using wings having advanced airfoil designs which call for blunt trailing edges. However, such designs have the inherent problem of increasing the base drag on the wing, thus comprising their benefits. The prior art tests of splitter plates do not suggest a solution.